Components of gas turbine engines are exposed to very high temperature, high pressure combustion gasses containing moisture, oxygen and other corrosive compounds. Modern gas turbine engines may have firing temperatures that exceed 1,400° C., and temperatures of 1,500–1,600° C. are expected as the demand for even more efficient engines continues. Cobalt and nickel base superalloys are used to form many gas turbine components, but even these superalloy materials must be aggressively cooled and/or insulated from the hot gas flow in order to survive long term operation in the combustion environment.
Ceramic matrix composite (CMC) materials have many potential applications in high temperature environments due to their ability to withstand and operate at temperatures in excess of those allowed for a non-insulated superalloy part. However, CMC's can survive temperatures in excess of 1,200° C. for only limited time periods in a combustion environment. Furthermore, oxide-based CMC's can not be cooled effectively with active cooling systems due to their low thermal conductivity and their limitations in cooling fluid path design due to manufacturing constraints. Non-oxide based CMCs can be aggressively cooled to withstand temperatures above 1200° C., but they are subject to environmental degradation that limits their useful life. To increase the operating temperature range and useful life for CMC materials, a high temperature insulation for a ceramic matrix composite material is described in U.S. Pat. No. 6,013,592.
Current structural ceramic technology for gas turbine engines relies on silica-based materials. Silica-based non-oxides such as silicon carbide (SiC) and silicon nitride (Si3N4) are subject to both oxidation and attack by high temperature, high pressure water vapor. In this dual degradation mechanism, the silicon carbide or silicon nitride is oxidized to form a thermally grown oxide (SiO2) layer. This oxide layer then reacts with the high temperature, high pressure water vapor to form a volatile hydroxide species [Si(OH)x] which is then lost to the environment. Thus, surface recession occurs in a continual process as the protective SiO2 layer volatizes and the base ceramic oxidizes to replenish the lost SiO2. This process is enhanced by the high velocity gas stream in a gas turbine environment. Accordingly, environmental barrier coatings (EBC) have been developed to protect silica-based non-oxide ceramics from the combustion environment. U.S. Pat. No. 5,391,404 describes a process for coating a silica-based ceramic with mullite, and U.S. Pat. No. 5,985,470 describes a barium strontium aluminosilicate (BSAS) bond coat underlying a thermally insulating top coat over a silicon carbide containing substrate. These EBC's typically function at a maximum surface temperature of 1,200–1,300° C. Since growth of a silicon dioxide layer underneath the environmental barrier coating could result in spalling of the coating and loss of environmental protection, the environmental barrier coating material must be sufficiently dense to prevent the ingress of oxygen through the coating, for example having only closed porosity of no more than approximately 10%.
The composite structure described in U.S. Pat. No. 6,013,592 utilizes a thick mullite-based thermal barrier coating over a ceramic matrix composite substrate material. Oxide ceramics such as mullite (3Al2O3–2SiO2) are not subject to oxidation, but they are degraded by the effects of high temperature water vapor, albeit at a slower rate than non-oxide ceramics. The rate of silica loss and subsequent recession of an oxide ceramic material will increase with an increasing temperature and flow velocity, and mullite may not perform adequately in certain gas turbine applications where flow rates are high and temperatures may be in the range of 1,500–1,600° C.